1. Field of the Invention
This invention relates to FLADE aircraft gas turbine engine and more particularly to construction and operating method of such an engine to reduce inlet spillage drag and reduce the infrared signature of the aircraft and aircraft engine.
2. Description of Related Art
A considerable effort has been made toward developing high performance variable cycle gas turbine engines. These types of engines have a unique ability to operate efficiently at various thrust settings and flight speeds both subsonic and supersonic. An important feature of the variable cycle gas turbine engine which contributes to its high performance is its capability of maintaining a substantially constant inlet airflow as its thrust is varied. This feature leads to important performance advantages under less than full power engine settings or maximum thrust conditions, such as during subsonic cruise. The effect of maintaining inlet airflow as thrust is reduced is to decrease such performance penalties as inlet spillage drag and afterbody closure drag, both of which have a considerable detrimental effect on the overall efficiency, size, weight, and performance of the aircraft and its engine.
Certain variable cycle engines, such as those described in U.S. Pat. Nos. 4,068,471 and 4,285,194, assigned to the same assignee as the present invention, achieve relatively constant airflow as thrust is varied by changing the amount of fan bypass flow with a valve system referred to as a variable area bypass injector (VABI). As engine thrust is decreased, the VABI increases bypass flow to offset decreasing core engine flow resulting in a relatively constant total engine flow, thus exhibiting the performance benefits described above. However, it uses very expensive multi-stage fan air to do so. Furthermore, it is somewhat limited in its ability to ingest the air at subsonic part power engine settings and engine operation because it has to be sized to match the flow conditions in the bypass duct into which is dumped some very highly pressurized air. This fan air is divided between the core flow and the bypass flow and therefore is controlled by the various bypass flow control mechanisms such as the VABI's. This in turn limits the degree to which the bypass air can be used to avoid spillage. Typically, conventional variable cycle engines have a limited range of thrust settings for a given Mach No. (particularly subsonic levels e.g. Mach No.=0.8-0.9) through which essentially constant airflow can be maintained and specific fuel consumption can be minimized. Therefore, it is desirable to construct and operate an aircraft gas turbine engine able to maintain inlet airflow at subsonic part power thrust settings more efficiently and over a broader flight envelope than is available in the prior art.
One particular type of variable cycle engine called a FLADE engine (FLADE being an acronym for "fan on blade") is characterized by an outer fan driven by a radially inner fan and discharging its flade air into an outer fan duct which is generally co-annular with and circumscribes an inner fan duct circumscribing the inner fan. One such engine, disclosed in U.S. Pat. No. 4,043,121, entitled "Two Spool Variable Cycle Engine", by Thomas et al., provides a flade fan and outer fan duct within which variable guide vanes control the cycle variability by controlling the amount of air passing through the flade outer fan duct.
There remains an important need to provide a high performance aircraft gas turbine engine, particularly of the variable cycle type, that is capable of maintaining an essentially constant inlet airflow over a relatively wide range of thrust at a given set of subsonic flight ambient conditions such as altitude and flight Mach No. in order to avoid spillage drag and to do so over a range of flight condition. This capability is particularly needed for subsonic part power engine operating conditions.
This invention also relates to an apparatus for suppressing and masking infrared (IR) emissions from engine exhaust ducts. The successful operation of combat aircraft is dependent, in part, upon the ability of the aircraft to remain undetected by infrared sensors of missiles during flight. The high temperatures of the engine's exhaust gases and the hot metal turbine parts and the hot metal walls directly in contact with the hot gases cause the engine to emit high levels of infrared energy. Military aircraft engaged in combat are vulnerable to anti-aircraft missiles employing highly sophisticated infrared sensors.
A number of apparatus have been designed to reduce infrared emissions from gas turbine engines. Each type of design endeavors to provide a combination of aerodynamics, heat transfer, and geometry which will result in an effective IR suppressor for the least suppressor weight and power effects on a turbine engine. One of these types of geometries utilizes a concentric center body within an annular duct. This suppressor geometry is referred to as a plug or center body suppressor and exemplified by U.S. Pat. Nos. 4,214,441, 4,044,555, 3,970,252 and the like. The plug suppressors are supported and fed cooling air from fan and/or high pressure bleed air aerodynamically shaped struts which are also used to position and support the center body. These hollow centerbody plug suppressors consume expensive fan and compressor air engine and power and result in reduced engine efficiency and combat operating radius.
The present invention provides a low power means of using a FLADE engine to improve overall engine efficiency and use the FLADE fan air as cooling air for IR suppression.